CN101275843A - Visual light imaging type autonomous navigation sensor system of middle and high orbit spacecraft - Google Patents
Visual light imaging type autonomous navigation sensor system of middle and high orbit spacecraft Download PDFInfo
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- CN101275843A CN101275843A CNA2007100910011A CN200710091001A CN101275843A CN 101275843 A CN101275843 A CN 101275843A CN A2007100910011 A CNA2007100910011 A CN A2007100910011A CN 200710091001 A CN200710091001 A CN 200710091001A CN 101275843 A CN101275843 A CN 101275843A
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Abstract
The invention provides a visible light imaging type auto-navigation sensor system for med-high orbit spacecraft, comprising an optical measurement imaging component, a detector focal plane component, a MEMS inertia measuring component, an information processing and error correction processing unit component. By using visible light spectrum segment detection, the invention can use normal optical glass to design, in order to reduce difficulty in developing optical system, remove the optical fiber converter and image intensifier, and reduce the apparatus complex, ensuring the detection accuracy via one-off imaging detection. The spectrum energy can be distributed by the filter in both of star and earth detection, thus, the spectrum energy distribution process can be carried out over identical optical system and detector. The system with integral design has advantages of light weight, small volume, low power consumption, high accuracy, high data updating rate and low costs.
Description
Technical field
The present invention relates to a kind of technology that is applied to spacecraft independent navigation attitude and orbit measurement system, specifically relate to a kind of photoelectronic imaging formula autonomous navigation sensor system of middle high orbit spacecraft.
Background technology
In spacecraft independent navigation field, there are multiple independent navigation attitude and positional information measuring system and method, disclose name as U.S. Honeywell Inc company in European Patent Publication No EP0589 387 A1 of application on September 20th, 1993 and be called " Method and System for Determining 3 AxisSpacecraft Attitude ", i.e. " three spacecraft attitudes are determined method and system ".Adopt the ultraviolet detector earth edge UV radiation profile of 280nm～300nm spectral coverage, determine the pitching and the roll attitude information in the earth's core, utilize same detector to survey and determine yaw-position information perpendicular to the fixed star direction vector of optical axis direction.System adopts refluxing reflection mirror compression visual field, adopts two hemisphere to add the optical fiber image rotator big visual field curved surface image planes are carried out imaging.Adopt data processor that the earth and the fixed star image information that collect are handled, obtain 3 attitude informations.Though this scheme has solved the problems of measurement of three-axis attitude and orbit altitude.But the deficiency that exists is that the optical system material of employing ultraviolet spectral coverage is less, adopts semiglobe lens and fibre optic image transmission too complicated, the cost height; Fibre optic image transmission and image intensifier bring additional noise in conjunction with meeting, reduce precision.
U.S. NASA has announced a project in the works in its new flourishing age, be referred to as " inertia star gyro " (Inertial Stellar Compass), adopt star sensor and MEMS gyro composite design, utilize the nearly drift of proofreading and correct gyro in real time of high-precision attitude information of star sensor.The deficiency of this scheme is, star sensor is single, can provide higher precision on optical axis direction, but on perpendicular to the direction of optical axis nearly 1 magnitude of precise decreasing, therefore the MEMS gyro drift correction accuracy for this direction just is affected.
" system emulation journal " in March, 2005 Vol.17, No3, the article that P529 delivers " makes up big visual field star sensor starlight refraction autonomous navigation of satellite method and emulation thereof ", and described sensor adopts 3 common star sensor space intersection hexagonal angles to constitute combined system, observe 3 fixed stars at earth edge simultaneously, release accurate the earth's core vector according to the atmospheric refraction model.The weak point of this scheme is to have adopted 3 star sensors, and cost is higher, makes the optical axis intersection of 3 star sensors adjust the done with high accuracy difficulty.
U.S. Microcosm company has developed a kind of autonomous navigation system MANS (MicrocosmAutonomous Navigation System), comprising earth sensor, the Sun and the Moon sensor, star sensor, gyro and accelerometer, owing to be that multi-sensor is united definite three-axis attitude and position, so precision is very high.But system is too complicated, and has adopted the double cone earth sensor that has movable part, cost height.
Content of the present invention
The objective of the invention is to overcome the shortcoming of above-mentioned prior art scheme, a kind of visual light imaging type autonomous navigation sensor system of middle high orbit spacecraft is proposed, it can use ordinary optical glass to design by adopting visible spectrum to survey, reduced the development difficulty of optical system, remove the complicacy that optic fiber converter and image intensifier have reduced instrument, adopted Polaroid detection to guarantee detection accuracy.Because of fixed star and survey of the earth are visible spectrum, distribute so can carry out spectral energy by means of optical filter, a shared optical system and detector have reduced cost.Determined that the earth's core vector and fixed star vector, two vectors determine that the three-axis attitude precision all can reach high precision owing to survey, so the same accuracy correction of drift of three on MEMS gyro.
The visual light imaging type autonomous navigation sensor system of middle high orbit spacecraft proposed by the invention, problems of measurement when the high orbit spacecraft does not rely on the integrated high-precision real of the three-axis attitude of satellite navigation system and orbit altitude in mainly solving.This sensor adopts the visible light detecting spectral coverage to overcome ultraviolet optics system complex and the high shortcoming of cost that U.S.'s ultraviolet sensors exists, overcome in the past the autonomous navigation sensor deficiency of scheme separately, such as the cost height that brings by distributing multi-sensor and complicated optical system, MEMS (micro electro mechanical system) gyro (Micro-electromechanical SystemsGyro is to call the MEMS gyro in the following text) by single optical sensor and three orthogonal directionss is inconsistent in conjunction with three precision that design brings, degenerate by the precision that image converter brings, the bulking value of being brought by multi-sensor distribution installation is big etc.
The objective of the invention is to realize by following technical proposals, the visual light imaging type autonomous navigation sensor system of high orbit spacecraft comprises optical measurement image-forming assembly, detector focal plane component, MEMS inertial measurement cluster, information processing and correction processing unit block in provided by the present invention, and described optical measurement image-forming assembly comprises imaging lens and beam-splitter structure.The photosurface of the detector of described detector focal plane component is installed on the imaging surface of optical measurement image-forming assembly, and the detector focal plane component will be fixed on the supporting construction of sensor system.Described MEMS inertial measurement cluster then comprises MEMS gyro and 3 accelerometers that quadrature is installed that 3 quadratures are installed, and each direction of principal axis of optical measurement coordinate system is parallel to MEMS gyro and 3 accelerometers that quadrature is installed that 3 quadratures are installed respectively.Described information processing and correction processing unit block are to adopt message handler that each sensor information is handled, and the zero shift that then the star sensor metrical information is used for the MEMS gyro is proofreaied and correct.Export nearly real-time high-precision three-axis attitude information and orbit altitude information by the normal data communication interface at last.
Star sensor in the visual light imaging type autonomous navigation sensor system of the middle high orbit spacecraft of integrated design and visible light quiescent imaging formula sensor are by adopting shared described optical imagery assembly of beam split optical filter and detector focal plane component, detector field of view is cut apart use, the central area is that visible light quiescent imaging formula sensor uses, be used for to earth imaging, fringe region is that star sensor uses, and is used for to the fixed star imaging.Described star sensor is a kind of by fixed star imaging extraction and standard star picture library coupling being obtained its optical axis with respect to the pointing vector of inertial space; Described visible light quiescent imaging formula sensor is a kind of by earth center vector and earth visual angle radius are extracted in earth imaging, utilizes the geometric relationship of earth visual angle radius and orbit altitude to determine orbit altitude simultaneously.Above-mentioned star sensor and visible light quiescent imaging formula sensor all have common image coordinates system, and its Z axle points to the ground direction of bowl along optical axis, and the row and column direction with detector array is consistent respectively with Y-axis for its X-axis.3 MEMS gyros all are the microelectromechanicpositioning gyros that adopts the MEMS technology to make, and they are installed in image coordinates respectively is on three parallel orthogonal axes directions of three axles.3 mems accelerometers all adopt the micro-electro-mechanical device of the measurement acceleration of motion of MEMS technology manufacturing, and weight is very light, and volume is very little, and three installation shaft are that three direction of principal axis are consistent with the star sensor image coordinates.Above MEMS gyro and mems accelerometer all belong to the MEMS inertial measurement cluster, and their installation shaft is that three change in coordinate axis direction are consistent with star sensor and visible light quiescent imaging formula earth sensor image coordinates, is beneficial to same reference measurement.
That system after the integrated design has is in light weight, volume is little, low in energy consumption, characteristics such as precision is high, data updating rate is high, cost is low.
Below just relevant technology contents of the present utility model and detailed description, existing conjunction with figs. and given embodiment describe as follows.
Description of drawings
Fig. 1 determines the autonomous navigation sensor structural representation of three-axis attitude and orbit altitude;
Fig. 2 is the autonomous navigation sensor optical measurement segmentation scheme schematic diagram of high orbit three-axis attitude in determining and orbit altitude;
Fig. 3 is the relation of optical measurement part image coordinates system with MEMS gyro and mems accelerometer measurement axis.
Embodiment
As Fig. 1-3, shown in, the visual light imaging type autonomous navigation sensor system of described middle high orbit spacecraft comprises optical measurement image-forming assembly 1, detector focal plane component 2, inertial measurement cluster 3, information processing and correction processing unit block 4.
Described optical measurement image-forming assembly 1, it comprises imaging lens and beam-splitter structure (not indicating).The photosurface of the detector of described detector focal plane component 2 is installed on the imaging surface of optical measurement image-forming assembly 1, and detector focal plane component 2 will be fixed on the supporting construction of sensor system.3 of described MEMS inertial measurement clusters comprise MEMS gyro and 3 accelerometers that quadrature is installed that 3 quadratures are installed, and each direction of principal axis of optical measurement coordinate system is parallel to the accelerometer that MEMS gyro that 3 quadratures install and 3 quadratures install (concrete mounting means referring to below in conjunction with the described content of Fig. 3) respectively.Described information processing and correction processing unit block 4 are to adopt message handler that each sensor information is handled, and the zero shift that then the star sensor metrical information is used for the MEMS gyro is proofreaied and correct.Export information such as nearly real-time high-precision three-axis attitude information and orbit altitude at last by the normal data communication interface.
Referring to Fig. 1, described optical measurement image-forming assembly 1 comprises imaging lens and beam-splitter structure (not marking).Described beam-splitter is divided into two passages with light path, one is Star Sensor imaging passage, another is an earth sensor imaging passage, the earth and the two brightness size before the track beam-splitter of fixed star are depended in see through and the selection of reflected light spectral coverage of beam-splitter, and the corresponding spectral coverage of establishing detector is from λ
1To λ
2, beam-splitter transmission spectral coverage is from λ
3To λ
4, transmitted light is from the radiation of the earth, because earth brightness is far longer than brightness, therefore from λ
3To λ
4Be at λ
1To λ
2Within narrow scope, the spectral transmittance of establishing optical measurement part is P (λ), the detector spectral response rate is K (λ), the brightness range of the earth before beam-splitter is L from weak to strong
E1～L
E2, the brightness range of the most weak fixed star before beam-splitter of detection is L from weak to strong
S1～L
S2, the dynamic range of detector is D, beam-splitter is Q for the reflectivity of stellar energy
1, beam-splitter is Q for the transmitance of earth energy
2, then have:
Q
2(L
e2+L
e1)≈?Q
1(L
s2+L
s1)..................................................(4)
Select to determine λ
3And λ
4, inequality (3) and approximate expression (4) are set up above making.
Sensor system as shown in Figure 1 also can be intercoursed the position with fixed star imaging passage and earth imaging passage, in this case, need the transmission spectral coverage of present beam-splitter change into the reflection and the reflection spectral coverage change transmission into.
3 the MEMS gyro measurement axis installation requirement separately of quadrature is parallel with detector image-forming coordinate system three axle x, y, z respectively each other, the zero shift error that they produce is separately proofreaied and correct by the star sensor metrical information respectively, and the kalman filter method of expansion is adopted in bearing calibration.Sensor will be exported the nearly angular speed and the attitude angle information in real time of 3 MEMS gyros, and error correction and information processing will be finished in information processing and correction processing unit block.
3 the mounting means of the mems accelerometer of quadrature is identical with 3 MEMS gyros each other, also is that 3 measurement axis are parallel to imaging coordinate system three axle x, y, z respectively.3 accelerometers are the instantaneous acceleration of measurement of x, y, three axles of z respectively, and twice integration obtains the displacement parameter of the relative initial position of satellite thus.More than calculate and in information processing and correction processing unit block, finish.
Information processing and correction processing unit block 4 are message handlers of sensor, being responsible for the star chart coupling of fixed star passage and the earth's core vector and the earth visual angle radius of earth passage extracts, and the fixed star vector that responsible Star Sensor is measured also is responsible for and communication of satellite control computer and multi-sensor Comprehensive Treatment for Information the zero shift correction of MEMS gyro and the integral operation of accelerometer.
This technical scheme combines optics attitude and earth visual angle radius measurement and inertia attitude and acceleration analysis, has unified measuring basis, has reduced measuring system ground systematic error; Simultaneously the drift of MEMS gyro is shifted near real-time correction and improved measuring accuracy.Can obtain high-precision fixed star vector and the earth's core vector by star sensor and earth sensor, therefore can obtain high-precision three-axis attitude measurement result, utilize the earth sensor passage can measure the visual angle radius of the earth simultaneously, can calculate out the flight track height by earth image extraction and optical performance parameter test result again, but they are discrete values.Having proofreaied and correct zero shift adopt the 3 axis MEMS gyro can obtain very high attitude and change resolution, but there is bigger null value drift in it, as long as therefore just can obtain high-precision MEMS gyro attitude measurement result.Because the MEMS gyro is measured the parallel installation of image coordinates axle with star sensor with earth sensor, therefore have and the same measuring basis of star sensor, by the drift that the high precision inertial space attitude of star sensor measurement can be proofreaied and correct the MEMS gyro well, this is characteristics of this programme.
Consider that the navigation measurement in full shadow district is subjected to the restriction of visible spectrum, earth sensor can not be worked, the less time interval of this section can adopt track extrapolation algorithm and the relative displacement of accelerometer measures satellite to change, and carries out the independent navigation based on star sensor, MEMS gyro, mems accelerometer.
Related invention scheme solved high precision that static independent navigation measures, near in real time, problem such as low-cost, complete autonomous, round-the-clock, have the following advantages:
(1) Star Sensor and earth sensor adopt visible spectrum to reduce the realization difficulty, have strengthened the sensor function;
(2) measurement target that adopts the optimized distribution method of inventing related beam-splitter spectrum transmitting section can take into account different brightness adopts same optical system and same detector image-forming.
(3) adopt star sensor, earth sensor, MEMS gyro, mems accelerometer to install and to reduce system errors for measurement, improve measuring accuracy with benchmark.
(4) adopt star sensor high-acruracy survey information to proofread and correct the zero shift of gyro at any time, can obtain nearly real-time high precision three-axis attitude information.
(5) adopt optics and the design of combined integratedization of inertia measurement can reduce dimensional weight and power consumption, multi-sensor information processing and correction processing can economize on resources, the advantage of performance information fusion.
Can finish full independent navigation measurement by round-the-clock, adopt star sensor, earth sensor, MEMS gyro and accelerometer can realize complete autonomous the measurement at sun according to the district, adopt star sensor, MEMS gyro and accelerometer can realize complete autonomous the measurement in the shadow region.
Referring again to Fig. 1, in the high orbit spacecraft photoelectronic imaging formula autonomous navigation sensor structure optical measurement image-forming assembly 1 its mainly to act on be by beam-splitter measuring system to be divided into two passages, these two passages can quadrature, also can be non-orthogonal, decide according to user demand.Determine the transmission and the reflected light spectral coverage of beam-splitter according to invention formula (3) and invention formula (4), make the earth and fixed star can be imaged on the same detector simultaneously, detector is selected the photoelectric detector of responding to visible light spectral coverage, as adopting CCD (Charge Coupled Devices, charge-coupled image sensor), also can adopt APS (Active PixelSensor, CMOS active pixel sensor).
Optical measurement image-forming assembly 1 is mainly for the earth and fixed star imaging, requirement has enough field angle, can beyond earth imaging viewing field, expand an annular visual field again, make fixed star imaging in annular visual field, the size of earth visual field and annular visual field is determined mainly once to catch fixed star imaging quantity on detector to be no less than 3 probability be standard greater than 99% so that the whole day ball is any.Mainly extract come out the earth's core vector and calculate earth visual angle radius of marginal information and match for earth image.Mainly extract asterism center of energy coordinate for the star chart picture and carry out star chart coupling extraction fixed star vector.
For the satellite of middle High Earth Orbit, because earth subtended angle is less, so imaging optical system adopts big visual field single lens to realize.
Referring to Fig. 3, MEMS inertial measurement cluster 3 comprises MEMS gyro and 3 accelerometers that quadrature is installed that 3 quadratures are installed, each direction of principal axis of optical measurement coordinate system is parallel to MEMS gyro and 3 accelerometers that quadrature is installed that 3 quadratures are installed respectively, as shown in Figure 3,31 is the detector image-forming face among the figure, and x, y, z are respectively detector image coordinates axle; 32 is that Star Sensor and earth sensor are optical system shared; 33 is 3 MEMS gyros of quadrature each other; 34 is 3 mems accelerometers of quadrature each other.3 MEMS gyros are installed in respectively in the plane parallel with xy, xz, yz, measurement axis x separately
1, y
1, z
1Parallel with respective x, y, z axle respectively; 3 mems accelerometers are installed in respectively in the plane parallel with xy, xz, yz, measurement axis x separately
2, y
2, z
2Parallel with respective x, y, z axle respectively.The installation site of each inertial measurement cluster can be adjusted under this condition.
Referring to Fig. 2, it is the schematic diagram of the photoelectronic imaging formula autonomous navigation sensor embodiment of middle high orbit spacecraft below.21 is imaging detector, is digital photoelectricity image device, as CCD (Charge CoupledDevices, charge-coupled image sensor) and APS (Active Pixel Sensor, CMOS active pixel sensor) etc.22 is Star Sensor and the shared optical system of earth sensor, adopts visible light design spectral coverage, adopts the single lens form for middle high orbit, and earth image is in the detector field of view central area, and the star chart picture is at the detector field of view fringe region; 23 is beam-splitter, and optical system is divided into Star Sensor passage and earth sensor passage, and plays balanced effect for the brightness of the earth and fixed star.24 is fixed star imaging annular visual field, shown in the shadow region of Fig. 2.5 is the earth, and 6 is fixed star, and 7 is the picture of fixed star on detector, and 8 is the picture of the earth on detector.
Constitute each functional module of foregoing invention, as quiescent imaging earth sensor, star sensor, MEMS gyro, mems accelerometer can based on information process unit individually or combination in any use, to satisfy different application targets.Can use separately as the star sensor assembly, also can unite use with quiescent imaging formula earth sensor, can also and quiescent imaging formula earth sensor, MEMS gyro, mems accelerometer thrin or two groups and use, the output corresponding information.Reduce when foregoing invention under the situation of assembly kind, the corresponding kind of non-common sparing of class component can take down.If when only needing the earth's core vector measurement, beam-splitter and dependency structure thereof can remove, and the star sensor relevant portion in the image processing software can remove, and the MEMS assembly all can remove.
The described system of foregoing invention is except and position definite in the attitude that orbits the earth are determined, the attitude of being diversion and the independent navigation that can also be applied to other celestial body respective rail height are measured.
Above-mentioned explanation only is embodiments of the invention, and is non-for limiting embodiments of the invention; All personages who is familiar with this skill, it complies with feature category of the present invention, other equivalence of having done changes or modifies, and selects or change of shape, increase and decrease functional module type and quantity etc. as size, material, all should be encompassed in the following institute of the present invention claim.
Claims (4)
1. the visual light imaging type autonomous navigation sensor system of high orbit spacecraft in a kind, it is characterized in that it comprises optical measurement image-forming assembly, detector focal plane component, MEMS inertial measurement cluster, information processing and correction processing unit block, described optical measurement image-forming assembly comprises imaging lens and beam-splitter structure; The photosurface of the detector of described detector focal plane component is installed on the imaging surface of optical measurement image-forming assembly, and the detector focal plane component will be fixed on the supporting construction of sensor system; Described MEMS inertial measurement cluster then comprises MEMS gyro and 3 accelerometers that quadrature is installed that 3 quadratures are installed, and each direction of principal axis of optical measurement coordinate system is parallel to MEMS gyro and 3 accelerometers that quadrature is installed that 3 quadratures are installed respectively; Described information processing and correction processing unit block are to adopt message handler that each sensor information is handled, the zero shift that then the star sensor metrical information is used for the MEMS gyro is proofreaied and correct, and exports nearly real-time high-precision three-axis attitude information and orbit altitude information by the normal data communication interface at last.
2. the visual light imaging type autonomous navigation sensor system of high orbit spacecraft in according to claim 1, it is characterized in that star sensor in the visual light imaging type autonomous navigation sensor system of high orbit spacecraft and visible light quiescent imaging formula sensor are by adopting shared described optical imagery assembly of beam split optical filter and detector focal plane component, detector field of view is cut apart use, the central area is that visible light quiescent imaging formula sensor uses, be used for to earth imaging, fringe region is that star sensor uses, and is used for to the fixed star imaging.Described star sensor is a kind of by fixed star imaging extraction and standard star picture library coupling being obtained its optical axis with respect to the pointing vector of inertial space; Described visible light quiescent imaging formula sensor is a kind of by earth center vector is extracted in earth imaging; Above-mentioned star sensor and visible light quiescent imaging formula sensor all have common image coordinates system, and its Z axle points to the ground direction of bowl along optical axis, and the row and column direction with detector array is consistent respectively with Y-axis for its X-axis.
3. the visual light imaging type autonomous navigation sensor system of high orbit spacecraft in according to claim 1 is characterized in that it is on three parallel orthogonal axes directions of three axles that described 3 MEMS gyros are installed in image coordinates respectively; Three installation shaft of 3 mems accelerometers are that three direction of principal axis are consistent with the star sensor image coordinates; Described MEMS gyro is that three change in coordinate axis direction are consistent with their installation shaft of mems accelerometer with star sensor and visible light quiescent imaging formula earth sensor image coordinates.
4. the visual light imaging type autonomous navigation sensor system of high orbit spacecraft in according to claim 1, each functional module that it is characterized in that forming system is respectively star sensor, visible light quiescent imaging formula sensor, MEMS gyro, mems accelerometer, they can both based on information process unit separately or the combination in any collocation use.
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Cited By (9)
| Publication number | Priority date | Publication date | Assignee | Title |
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| CN101915927A (en) * | 2010-04-14 | 2010-12-15 | 清华大学 | Infrared measurement based system and method thereof for determining relative state of inner satellite |
| CN102175259A (en) * | 2010-12-31 | 2011-09-07 | 北京控制工程研究所 | Autonomous navigation simulation test system based on earth-sun-moon integrated sensor |
| CN102116641B (en) * | 2009-12-31 | 2012-08-08 | 北京控制工程研究所 | Semi-physical simulation testing system and method for deep space autonomous navigation star sensor |
| CN104764452A (en) * | 2015-04-23 | 2015-07-08 | 北京理工大学 | Hybrid position-posture tracking method based on inertia and optical tracking systems |
| CN105928526A (en) * | 2016-04-25 | 2016-09-07 | 航天东方红卫星有限公司 | Method for determining satellite attitude based on visible light earth sensor |
| CN107329191A (en) * | 2017-05-17 | 2017-11-07 | 上海卫星工程研究所 | GMS image-guidance and registering imaging test system and test method |
| CN107367275A (en) * | 2017-06-19 | 2017-11-21 | 北京控制工程研究所 | A kind of high rail autonomous navigation of satellite sensor |
| CN111707291A (en) * | 2020-06-23 | 2020-09-25 | 上海航天控制技术研究所 | Automatic assembling and calibrating device and automatic assembling and calibrating method for star sensor focal plane |
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Family Cites Families (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US5396326A (en) * | 1989-04-03 | 1995-03-07 | Northrop Grumman Corporation | Two gimbal error averaging astro-inertial navigator |
| US5442560A (en) * | 1993-07-29 | 1995-08-15 | Honeywell, Inc. | Integrated guidance system and method for providing guidance to a projectile on a trajectory |
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- 2007-03-29 CN CN2007100910011A patent/CN101275843B/en not_active Expired - Fee Related
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| CN102116641B (en) * | 2009-12-31 | 2012-08-08 | 北京控制工程研究所 | Semi-physical simulation testing system and method for deep space autonomous navigation star sensor |
| CN101915927A (en) * | 2010-04-14 | 2010-12-15 | 清华大学 | Infrared measurement based system and method thereof for determining relative state of inner satellite |
| CN102175259A (en) * | 2010-12-31 | 2011-09-07 | 北京控制工程研究所 | Autonomous navigation simulation test system based on earth-sun-moon integrated sensor |
| CN102175259B (en) * | 2010-12-31 | 2012-11-14 | 北京控制工程研究所 | Autonomous navigation simulation test system based on earth-sun-moon integrated sensor |
| CN104764452A (en) * | 2015-04-23 | 2015-07-08 | 北京理工大学 | Hybrid position-posture tracking method based on inertia and optical tracking systems |
| CN105928526A (en) * | 2016-04-25 | 2016-09-07 | 航天东方红卫星有限公司 | Method for determining satellite attitude based on visible light earth sensor |
| CN107329191A (en) * | 2017-05-17 | 2017-11-07 | 上海卫星工程研究所 | GMS image-guidance and registering imaging test system and test method |
| CN107329191B (en) * | 2017-05-17 | 2020-04-21 | 上海卫星工程研究所 | Imaging test system and method for image navigation and registration of static meteorological satellite |
| CN107367275A (en) * | 2017-06-19 | 2017-11-21 | 北京控制工程研究所 | A kind of high rail autonomous navigation of satellite sensor |
| CN111707291A (en) * | 2020-06-23 | 2020-09-25 | 上海航天控制技术研究所 | Automatic assembling and calibrating device and automatic assembling and calibrating method for star sensor focal plane |
| CN114142926A (en) * | 2021-10-15 | 2022-03-04 | 北京遥测技术研究所 | Ultra-far deep space laser communication capturing and tracking system |
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